This invention relates generally to aircraft gas turbine engines and, more particularly, to a structural cooling air manifold associated with a high temperature, variable area, low pressure turbine.
It is by now understood that in gas turbine engines energy is added to the air through the processes of compression and combustion, while energy is extracted by means of a turbine. In a turbofan engine, compression is accomplished sequentially through a fan and thereafter through a multistage compressor, the fan and compressor being independently driven by a high pressure and a low pressure turbine, respectively, through concentric shaft connections. Combustion occurs between the multistage compressor and the high pressure turbine. Since the energy available to the turbines far exceeds that required to maintain the compression process, the excess energy is exahausted as high velocity gases through one or more nozzles at the rear of the engine to produce thrust by the reaction principle.
Since the fan and compressor are on separate concentric shafts and are driven by separate, axially spaced turbines, a means for regulating their relative rotational speeds is desirable for performance optimization. Further, it becomes desirable to control the relative amounts of energy added by the fan and compressor which, in turn, are controlled by how much energy is extracted by their respective turbines. It can be appreciated that the faster the fan or compressor rotates, the more air it pumps, and vice versa. Furthermore, it is recognized that if a stage of stationary turbine vanes may be made to provide a variable flow area through the turbine by making the vanes rotatable about their respective longitudinal axes, the energy extraction characteristics of either of the high or low pressure turbines may be modulated. Thus, if the capability of the high pressure turbine to extract energy was reduced, more energy would be available to the low pressure turbine and the fan could be driven at a higher rotational speed relative to the compressor, and vice versa. This ability to regulate the relationship between fan and compressor rotational speeds is extremely important in designing the most efficient engine over a range of operating conditions. Such optimized engines have recently been referred to as variable cycle engines and are characterized as possessing variable geometry components in order to optimize performance for both subsonic and supersonic cruise, for example. It is characteristic of some of these variable cycle engines that both the high and low pressure turbines are of the variable area variety for maximum modulation of energy extraction.
Additionally, it is well understood that gas turbine engine shaft horsepower and specific fuel consumption, which is the rate of fuel consumption per unit of power output, can be improved by increasing turbine inlet temperatures to provide more energy for extraction. To permit turbines to operate at gas stream temperatures which are higher than the materials can normally tolerate, and to take advantage of the potential performance improvements associated with higher turbine inlet temperatures, considerable effort has been devoted to the development of sophisticated methods of turbine cooling. Modern cooling technology utilizes hollow turbine vanes and blades to permit operation at turbine inlet gas temperatures well in excess of 2300.degree. F. (1260.degree. C.). Various techniques have been devised to cool these hollow blades and vanes incorporating convection, impingement and film cooling, either singly or in combination. U.S. Pat. Nos. 3,700,348 and 3,715,170, assigned to the assignee of the present invention, are excellent examples of advanced turbine air cooling technology incorporating these basic cooling concepts.
However, air cooling has generally been limited in application to the high pressure turbine which is exposed to the highest combustion temperatures, and the delivery of cooling air to the high pressure turbine components has been relatively straightforward because of the proximity of these parts to the multistage compressor from which the relatively cool air is extracted as a source of coolant. But, in advanced high temperature turbofans and variable cycle engines, the low pressure turbine often requires the same considerations with regard to cooling as contemporary high pressure turbines. Unfortunately, these turbines do not enjoy the proximity to the coolant air source as the high pressure turbine. Furthermore, if the low pressure turbine is of the variable area variety in order to enhance energy modulation, the situation is compounded in that coolant air must be provided to the interior of movable vanes. Therefore, a structure is required which can not only support a stage of variable area vanes to provide efficient operation over a range of operating conditions, but which can also ensure that the vanes receive an adequate supply of cooling air in order to take advantage of the potential performance improvements associated with higher low pressure turbine inlet temperatures. It is also necessary that such a structure is of the lightest possible weight consistent with modern aircraft technology.